The present invention relates to a cooling system for a stationary vane in a gas turbine. More specifically, the present invention relates to a cooling air system in which compressed air is recirculated through the vane.
A gas turbine is comprised of a compressor section that produces compressed air that is subsequently heated by burning fuel in a combustion section. The hot gas from the combustion section is directed to a turbine section where the hot gas is used to drive a rotor shaft to produce power. The combustion section is typically comprised of a shell that forms a chamber that receives compressed air from the compressor section. A plurality of cylindrical combustors are disposed in the chamber and receive the compressed air along with the fuel to be burned. A duct is connected to the aft end of each combustor and serves to direct the hot gas from the combustor to the turbine section.
The turbine section typically employs a plurality of stationary vanes circumferentially arranged in rows. Since such vanes are exposed to the hot gas discharging from the combustion section, cooling of these vanes is of utmost importance. Traditionally, cooling was accomplished by flowing cooling air from the chamber through a cavity formed in the airfoil portion of the vane, which is essentially hollow. Typically, a number of small passages are formed inside the vane airfoil that extend from the cavity to the surfaces of the vane, such as the leading and trailing edges or the suction and pressure surfaces. Often, such as in the case of leading edge cooling, the passages direct the cooling air from the cavity so that it flows over the surface of the vane in a thin film, thereby cooling the vane in what is often referred to as "film cooling." In any case, after the cooling air exits the vane passages, it enters and mixes with the hot gas flowing through the turbine section.
Unfortunately, the traditional approach to cooling the turbine vanes has a detrimental impact on the thermal efficiency of the gas turbine. Although the cooling air eventually mixes with the hot gas expanding in the turbine, since it bypasses the combustion process the work recovered from the expansion of the compressed cooling air is much less than that recovered from the expansion of the compressed air heated in the combustors. In fact, as a result of losses due to pressure drop and mechanical efficiency, the work recovered from the cooling air is less than that required to compress the air in the compressor. In addition, discharging the cooling air into the hot gas flow results in aerodynamic losses as the cooling air mixes with the hot gas.
Moreover, although the traditional scheme may provide adequate cooling of the airfoil portion of the vane, albeit at the expense of efficiency, it does not provide thorough cooling of the inner and outer shroud portions of the vane, so that deterioration of these components can limit the operating temperature of the turbine.
It is therefore desirable to provide a cooling system for a gas turbine vane that does not result in discharging cooling air directly into the hot gas flow path by bypassing the combustors, and that achieves effective cooling of the inner and outer shrouds.